CN115324731A

CN115324731A - Gas turbine

Info

Publication number: CN115324731A
Application number: CN202210983590.9A
Authority: CN
Inventors: 段萌珠
Original assignee: Xingchen Mengxiang Technology Beijing Co ltd
Current assignee: Xingchen Mengxiang Technology Beijing Co ltd
Priority date: 2022-08-16
Filing date: 2022-08-16
Publication date: 2022-11-11

Abstract

The invention discloses a gas turbine, which comprises a first rotating shaft, a second rotating shaft and a combustion chamber, wherein the first rotating shaft and the second rotating shaft are sequentially arranged along the axial direction, one end of the first rotating shaft, which is far away from the second rotating shaft along the axial direction, is provided with a first gas compressor, one end of the first rotating shaft, which is close to the second rotating shaft along the axial direction, is provided with a first turbine, one end of the second rotating shaft, which is far away from the first rotating shaft along the axial direction, is provided with a second gas compressor, one end of the second rotating shaft, which is close to the first rotating shaft along the axial direction, is provided with a second turbine and a third turbine sequentially, the gas inlet end of the second gas compressor is communicated with the gas outlet end of the first gas compressor, the gas outlet end of the second gas compressor is communicated with the gas inlet end of the combustion chamber, the gas outlet end of the combustion chamber is communicated with the gas inlet end of the third turbine, and the third turbine is a gas film cooling turbine with a hollow structure. The gas turbine disclosed by the invention has the advantages that the gas compressor and the turbine are reasonably arranged, the cooling of the turbine can be enhanced, and further, the temperature of the inlet of the turbine can be further improved, so that the efficiency of the gas turbine is improved.

Description

Gas turbine

Technical Field

The invention belongs to the field of heat engines, and particularly relates to a gas turbine.

Background

The gas turbine mainly comprises three parts of a gas compressor, a combustion chamber and a turbine, is matched with an air inlet system, an air exhaust system, a control system, a transmission system and other auxiliary systems, takes air as a medium, and is a rotary power machine for converting heat energy generated by fuel combustion into mechanical work and outputting the mechanical work. The working process is as follows: the compressor driven by the turbine to rotate continuously sucks air from the atmosphere and compresses and boosts the air, the compressed air enters the combustion chamber and is mixed and combusted with the injected fuel to become high-temperature gas, the high-temperature gas flows into the turbine to expand and do work, and the pressure of the gas after doing work is reduced to the atmospheric pressure and is finally discharged into the atmosphere. The high-temperature gas formed after combustion heating and temperature rise has greatly improved work-doing capability, so that the work output of the turbine is obviously greater than the power consumption of the gas compressor, and more surplus work is output externally to drive the load.

It has been shown that the thermal efficiency and power output of a gas turbine increase with increasing turbine inlet temperature, typically 10% for every 40 ℃ increase in turbine inlet temperature and 1.5% for thermal efficiency [1]. However, the increase of the turbine inlet temperature is limited by the temperature tolerance of the turbine blade material directly exposed to the high-temperature combustion gas and the cooling effect of the turbine blade, and the great increase of the turbine inlet temperature may cause the structural strength of the turbine blade to be reduced, thereby causing the deformation or ablation failure of the blade. Therefore, the turbine blade directly determines the performance level of the gas turbine to a certain extent, and the key technical problem of developing a high-performance gas turbine is how to improve the turbine inlet temperature on the premise of meeting the requirement of long-service-life safe operation of the gas turbine.

Li laugh hall, hou ling cloud, yanmin, etc. modern gas turbine technology [ M ]. Beijing: aeronautical industry press, 2006.

Disclosure of Invention

The invention aims to provide a gas turbine aiming at the defects in the prior art, wherein a gas compressor and a turbine are reasonably arranged, the cooling of the turbine can be enhanced, and the temperature of the inlet of the turbine can be further increased so as to improve the efficiency of the gas turbine.

The invention adopts the following technical scheme:

in a first aspect, an embodiment of the present invention provides a gas turbine, including a first rotating shaft, a second rotating shaft, and a combustion chamber, which are sequentially arranged along an axial direction, wherein a first gas compressor is installed at one end of the first rotating shaft, which is far away from the second rotating shaft along the axial direction, and a first turbine is installed at one end of the first rotating shaft, which is close to the second rotating shaft along the axial direction, a second gas compressor is installed at one end of the second rotating shaft, which is far away from the first rotating shaft along the axial direction, and a second turbine and a third turbine are sequentially installed at one end of the second rotating shaft, which is close to the first rotating shaft along the axial direction, an air inlet end of the second gas compressor is communicated with an air outlet end of the first gas compressor, an air outlet end of the second gas compressor is communicated with an air inlet end of the combustion chamber, an air outlet end of the combustion chamber is communicated with an inlet end of the third turbine, and the third turbine is a gas film cooling turbine with a hollow structure.

Furthermore, at least one axial end of the second rotating shaft is provided with an at least partially built-in inner air bearing, the inner air bearing is provided with an air supply channel, and the air supply channel supplies cooling air for the third turbine through the inside of the second rotating shaft.

Furthermore, one end of the second rotating shaft, which is provided with the inner air bearing, is provided with a nose cone, and the inner air bearing is provided with a nose cone air supply hole.

Furthermore, the gas turbine also comprises a plurality of first radial bearings which are externally arranged radial air bearings sleeved on the outer peripheral surfaces of the first rotating shaft and/or the second rotating shaft.

Further, the second turbine is located between the first turbine and the third turbine in the axial direction, and the gas from the exhaust end of the combustion chamber passes through the third turbine, the second turbine and the first turbine in sequence and then is discharged.

Further, the gas turbine also comprises a first stationary blade assembly, the first stationary blade assembly is located between a third turbine and the exhaust end of the combustion chamber and is arranged close to the third turbine, and the stationary blade in the first stationary blade assembly is a fluid cooling blade with a hollow structure.

Further, the second turbine is a film-cooled turbine having a hollow structure.

Further, the gas turbine further includes a second vane assembly located between and adjacent to the third turbine and the second turbine, the vanes in the second vane assembly being fluid-cooled blades having a hollow structure.

The gas turbine further comprises a first motor and a second motor, wherein the first motor is arranged on the first rotating shaft and is positioned between the first gas compressor and the first turbine or positioned on one side of the first gas compressor, which is far away from the first turbine; the second motor is arranged on the second rotating shaft and is positioned on one side, far away from the third turbine, of the second compressor.

Further, the power of the first motor is greater than or equal to the power of the second motor.

Further, the second rotating shaft comprises a first shaft section and a second shaft section which are arranged along the axial direction, the first shaft section and the second shaft section are connected through a coupler, and the coupler is provided with an air seal cavity communicated with air passages in the first shaft section and the second shaft section.

Further, a second motor is mounted to the second shaft section.

Further, the gas turbine includes an inner air bearing located at an end of the second shaft section axially remote from the first shaft section.

Further, the first electric machine and/or the second electric machine is a superconducting electric machine, and the fuel of the gas turbine is reused as a coolant for the superconducting electric machine.

Further, the pressure ratio of the second compressor is greater than or equal to that of the first compressor.

Further, the second compressor is a centrifugal compressor.

Further, the first compressor is selected from one or more of a centrifugal compressor, an axial compressor and a diagonal compressor.

Further, the gas turbine also comprises a heat regenerator HR, the heat regenerator HR comprises a first passage and a second passage which can exchange heat, the exhaust end of the second compressor is communicated with the air inlet end of the first passage, the air inlet end of the combustion chamber is communicated with the exhaust end of the first passage, the exhaust end of the first turbine is communicated with the air inlet end of the second passage, and the exhaust end of the second passage is communicated with the external atmosphere or other external components.

According to the gas turbine provided by the embodiment of the invention, the first compressor is arranged on the first rotating shaft, the second compressor is arranged on the second rotating shaft, and the air inlet end of the second compressor is communicated with the air outlet end of the first compressor, so that the pressure ratio of the compressors is improved, the efficiency of the gas turbine is favorably improved, and the third turbine communicated with the air outlet end of the combustion chamber is set as an air film cooling turbine with a hollow structure, and the cooling air film protects the turbine, so that the gas temperature before the turbine can be further improved, and the efficiency of the gas turbine is further improved. And the first turbine, the second turbine and the third turbine are arranged on the first rotating shaft and the second rotating shaft, so that the high-temperature gas with the increased temperature can be more fully utilized, and the overall efficiency of the gas turbine is improved.

Drawings

FIG. 1 is a schematic block diagram of a gas turbine according to an embodiment of the present invention;

FIG. 2 is a partial schematic illustration of an embodiment of the third turbine of FIG. 1;

FIG. 3 is a partial schematic illustration of another embodiment of the third turbine of FIG. 1;

FIG. 4 is a partial schematic illustration of yet another embodiment of the third turbine of FIG. 1;

FIG. 5 is a partial schematic view of one embodiment of the second shaft of FIG. 1;

FIG. 6 is a partial schematic view of an embodiment of the inner air bearing of FIG. 1;

FIG. 7 is a partial schematic view of another embodiment of the inner air bearing of FIG. 1;

FIG. 8 is a partial schematic illustration of yet another embodiment of the third turbine of FIG. 1;

FIG. 9 is a schematic block diagram of a gas turbine according to another embodiment of the present invention;

FIG. 10 is a schematic block diagram of a gas turbine according to yet another embodiment of the present invention;

FIG. 11 is a schematic view of an embodiment of the second shaft of FIG. 10;

FIG. 12 is a schematic block diagram of a gas turbine according to yet another embodiment of the present invention;

FIG. 13 is a schematic block diagram of a gas turbine according to yet another embodiment of the present invention;

FIG. 14 is a schematic block diagram of a gas turbine according to yet another embodiment of the present invention.

Reference numerals:

Detailed Description

Features and exemplary embodiments of various aspects of the present invention will be described in detail below, and in order to make objects, technical solutions and advantages of the present invention more apparent, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. It should be understood that the specific embodiments described herein are merely illustrative of the invention and are not to be construed as limiting the invention. It will be apparent to one skilled in the art that the present invention may be practiced without some of these specific details. The following description of the embodiments is merely intended to provide a better understanding of the present invention by illustrating examples of the present invention.

The efficiency of a gas turbine is positively correlated to the pre-turbine temperature. However, when the temperature before the turbine is too high and exceeds the heat resistance limit of the turbine material, the turbine cannot normally and continuously operate, and therefore, efforts are made to develop a material having a higher heat resistance temperature, but the development progress is slow and the price of a new material is high, resulting in a great increase in the cost of the gas turbine.

The inventors have discovered that the turbine may be protected by film or liquid cooling to improve the turbine's material stability at higher pre-turbine temperatures. However, the centrifugal force of the rotation of the turbine alone is used to increase the cooling air pressure, which is often lower or slightly greater than the static pressure in front of the turbine, and thus an effective cooling air film is not formed. The liquid film cooling requires a high cooling liquid, such as deionized water, and the cooling liquid is consumed continuously, so that the cooling liquid is difficult to recycle, and the cost is increased.

Therefore, there is a need for a gas turbine that can improve the cooling effect of the turbine, and thus can achieve an improvement in efficiency by increasing the temperature in front of the turbine, and to ensure that cooling is continuous and easy to achieve, and to control manufacturing and modification costs. Meanwhile, the gas compressor and the turbine need to be reasonably arranged so as to fully utilize the high-temperature gas with the increased temperature.

To achieve the above object, fig. 1 is a schematic structural view of a gas turbine according to an embodiment of the present invention, and the present invention provides a gas turbine including a first rotating shaft 110 and a second rotating shaft 120, which are sequentially arranged in an axial direction, and a combustion chamber 400, as shown in fig. 1.

The first rotating shaft 110 and the second rotating shaft 120 may be made of steel, or may be made of other suitable metals, alloys, or composite materials. The first and

second shafts

110, 120 are supported by bearings to the casing or bearing housing of the gas turbine. The bearing is preferably an air bearing, or may be other non-contact bearing such as a magnetic bearing or a gas-magnetic hybrid bearing, or some of the bearings may be contact bearings such as ball bearings or roller bearings.

The first rotating shaft 110 and the second rotating shaft 120 may be coaxially arranged, and an axial end of the first rotating shaft 110 is opposite to an axial end of the second rotating shaft 120.

The first compressor 210 is installed at one end of the first rotating shaft 110, which is far away from the second rotating shaft 120 in the axial direction, and the first turbine 310 is installed at one end of the first rotating shaft 110, which is close to the second rotating shaft 120 in the axial direction.

The first compressor 210 is selected from one or more of a centrifugal compressor, an axial compressor, and a diagonal compressor. The first compressor 210 may be a single stage centrifugal compressor, a multi-stage axial compressor (e.g., 3 or 4 stages), a multi-stage diagonal compressor (e.g., 3 stages), or a combination of a centrifugal compressor, an axial compressor, and a diagonal compressor. In some embodiments, the first compressor 210 may include a compressor wheel and a diffuser. The intake end of the first compressor 210 communicates with the outside environment for air induction.

The second compressor 220 is installed at one end of the second rotating shaft 120 that is far away from the first rotating shaft 110 along the axial direction, and the second turbine 320 and the third turbine 330 are sequentially installed at one end of the second rotating shaft that is close to the first rotating shaft 110 along the axial direction.

The second compressor 220 may be a centrifugal compressor. In some embodiments, the second compressor 220 may include a compressor wheel and a diffuser. An air inlet end of the second compressor 220 is communicated with an air outlet end of the first compressor 210, and an air outlet end of the second compressor 220 is communicated with an air inlet end of the combustion chamber 400. The first compressor 210 is connected in series with the second compressor 220, and ambient external air (e.g., air) is compressed by the first compressor 210 and then enters the second compressor 220 for further compression to increase the final pressure of the air leaving the compressor combination. The combined pressure ratio of the first compressor 210 and the second compressor 220 connected in series in the above manner is greater than the pressure ratio of the first compressor 210 alone or the second compressor 220 alone. The pressure ratio of the first compressor 210 may be selected from 2 to 5, and the pressure ratio of the second compressor 220 may be selected from 5 to 10. For example, if the pressure ratio of the first compressor 210 is 5 and the pressure ratio of the second compressor 220 is 5, the pressure ratio after series connection is 25; if the pressure ratio of the first compressor 210 is 3 and the pressure ratio of the second compressor 220 is 8, the pressure ratio after series connection is 24; if the pressure ratio of the first compressor 210 is 4 and the pressure ratio of the second compressor 220 is 7, the tandem back pressure ratio is 28. In some alternative embodiments, the pressure ratio of the second compressor 220 is equal to or greater than the pressure ratio of the first compressor 210.

The combustor 400 may be an annular combustor, a mono-can combustor, a can-annular combustor, or the like. When the combustion chamber 400 is an annular combustion chamber, the combustion chamber 400 may be disposed about the second axis of rotation 120. The exhaust end of the combustor 400 communicates with the inlet end of the third turbine 330. In the axial direction, the second turbine 320 is located between the first turbine 310 and the third turbine 330, and the combustion gas from the exhaust end of the combustor 400 passes through the third turbine 330, the second turbine 320 and the first turbine 310 in sequence and is then discharged.

The third turbine 330 is a film-cooled turbine having a hollow structure.

According to the gas turbine of the embodiment of the invention, the first compressor 210 is arranged on the first rotating shaft 110, the second compressor 220 is arranged on the second rotating shaft 120, and the air inlet end of the second compressor 220 is communicated with the air outlet end of the first compressor 210, so that the compressor pressure ratio after combination and series connection is improved, the efficiency of the gas turbine is improved, and the third turbine 330 communicated with the air outlet end of the combustion chamber 400 is set as a film cooling turbine with a hollow structure, and the cooling film is used for protecting the turbine 330, so that the gas temperature before the turbine can be further improved, and the efficiency of the gas turbine is improved. And by arranging the first turbine 310, the second turbine 320 and the third turbine 330 on the first rotating shaft 110 and the second rotating shaft 120, the high-temperature combustion gas with increased temperature can be more fully utilized to improve the overall efficiency of the gas turbine.

Further, as shown in fig. 2, fig. 2 is a partial schematic view of an embodiment of the third turbine in fig. 1, and fig. 2 exemplarily shows that the third turbine 330 is a hollow turbine, and the third turbine 330 includes a turbine housing 333 and a turbine inner cavity 334 surrounded by the turbine housing 333. The turbine housing 333 is provided with a turbine film hole 332, and the turbine film hole 332 is communicated with a turbine inner cavity 334. The hollow turbine housing 333 may be fixedly attached (e.g., welded) to two or more pieces of housing. In other embodiments, the turbine cavity 334 may also be a communication channel to supply air to the turbine film holes 332. In other embodiments, baffles, struts, or labyrinths may be provided within the turbine interior 334 to increase the cooling effect within the turbine to further enhance the overall cooling effect of the third turbine 330.

Referring to fig. 3, which is a partial schematic view of another embodiment of the third turbine shown in fig. 1, the number of the turbine film holes 332 is multiple, and the turbine film holes 332 may be located at the leading edges and the tips of the blades of the turbine 330. A plurality of turbine film holes 332 at the leading edge of the blade may be arranged in a radial direction of the turbine 330. The turbine film holes 332 are located at various locations to meet the cooling requirements at each location and to provide fine control over the formation of the cooling film. The turbine film holes 332 at the tip of the blade may be multiplexed as turbine-at-bearing jet holes. On one hand, the turbine film holes 332 on the blade tip are used for cooling the blade tip, on the other hand, radial gas thrust can be formed between the blade tip with the turbine film holes 332 and the outer shell, and the pressure gas sprayed out of the turbine film holes 332 on the blade tip can radially support the turbine 330 to a certain extent, so that the improvement of the rotational stability of the rotor system is facilitated.

As shown in FIG. 4, which is a partial schematic view of another embodiment of the third turbine shown in FIG. 1, the turbine film holes 332 may also be located at the trailing edges of the blades of the turbine 330, and the flow of the gas stream is facilitated by the exit from the trailing edges.

Further, as shown in fig. 5, fig. 5 is a partial schematic view of an embodiment of the second rotating shaft in fig. 1, and the second rotating shaft 120 has a shaft connecting through hole 123 at a position corresponding to the third turbine 330, so as to provide cooling air for the third turbine 330 through the shaft connecting through hole 123 through the inside of the second rotating shaft 120. Also, at least one axial end of the second rotating shaft 120 has an inner shaft bearing cavity 122, fig. 1 shows that the inner shaft bearing cavity 122 is located at an end far away from the first rotating shaft 110 in the axial direction, and in other embodiments, the inner shaft bearing cavity 122 may also be located at an end near to the first rotating shaft 110 in the axial direction. The shaft inside bearing chamber 122 communicates with the shaft connecting through hole 123 through the shaft inside air passage 124, and the shaft connecting through hole 123 communicates with the turbine chamber 334 through the turbine communicating hole 331 of the third turbine 330. Thus, the cooling air from the outside enters the shaft inner bearing cavity 122, then enters the turbine inner cavity 334 through the shaft inner air passage 124 and the shaft connecting through hole 123, and then is ejected through the turbine film hole 332 to form a cooling film on the outer wall surface of the third turbine 330.

Because the third turbine 330 is disposed adjacent to the exhaust end of the combustion chamber 400, the temperature of the exhaust gas of the combustion chamber 400 is highest at the position (R1) before the third turbine 330, the temperature at the position (R2) after passing through the third turbine 330 is reduced, the temperature at the position (R3) after passing through the second turbine 320 is further reduced, and the temperature at the position (R4) after passing through the first turbine 310 is the waste heat temperature, and at this time, the exhaust gas with waste heat can be directly discharged into the atmosphere or guided to other external components for further use. Therefore, of the first turbine 310, the second turbine 320, and the third turbine 330, the third turbine 330 has the highest pre-turbine temperature, so the third turbine 330 employs a film cooling turbine having a hollow structure to improve the cooling effect. Accordingly, the second turbine 320 may employ a film-cooled turbine having a hollow structure to enhance the cooling effect, or the second turbine 320 may be a conventional solid turbine when the pre-turbine temperature of the second turbine 320 is within a turbine material tolerance range. The first turbine 310 may generally be a conventional solid turbine.

Further, as shown in fig. 1, at least one axial end of the second rotating shaft 120 is provided with an at least partially built-in inner air bearing 510, and the inner air bearing 510 has an air supply passage for supplying cooling air to the third turbine 330 through the inside of the second rotating shaft 120.

In particular, with respect to the second shaft 120, the inner air bearing 510 is a stator component, and the inner air bearing 510 may be connected to a housing or a bearing housing of a gas turbine, for example. Inner air bearing 510 extends at least partially into shaft inner bearing cavity 122, and a gas film gap of the gas bearing is formed between the portion of inner air bearing 510 extending into shaft inner bearing cavity 122 and the circumferential inner wall of shaft inner bearing cavity 122. The inner air bearing 510 supports the second shaft 120 to rotate, and the inner air bearing 510 is located at one end of the second shaft 120 to at least partially close the opening of the shaft inner bearing cavity 122, and the remaining gap is air-sealed by an air film formed during the operation of the inner air bearing 510. Inner air bearing 510 has an air supply passage 511 in communication with shaft inner bearing cavity 122 to provide turbine cooling air to shaft inner bearing cavity 122.

Further, as shown in fig. 6, fig. 6 is a partial schematic view of an embodiment of the inner air bearing in fig. 1, the inner air bearing 510 has a plurality of first air holes 512 extending radially and arranged circumferentially, air inlets of the first air holes 512 are connected to an air source, and air outlets of the first air holes 512 face a circumferential wall surface of the shaft inner bearing cavity 122. The gas source may be an external gas source, such as a gas pump. The air supply may also be the second compressor 220, for example, the air inlet of the first air hole 512 is communicated with the air outlet end of the second compressor 220. In fig. 6, the inlet of the first air hole 512 is shown to be connected to the air supply passage 511, and the air supply passage 511 is connected to the air source, in other alternative embodiments, the inlet of the first air hole 512 may be connected to the air source through other passages.

Specifically, each first air hole 512 is located between the outer opening of the shaft inner bearing cavity 122 and the air outlet of the air supply passage 511 in the axial direction, so that the air outlet of the air supply passage 511 is located inside the shaft inner bearing cavity 122 and air sealing can be achieved through each first air hole 512.

During operation of the gas turbine, the first air holes 512 inject air toward the circumferential wall surface of the shaft inner bearing cavity 122 to form an air film, the air film supports the second rotating shaft 120 to rotate and forms an air seal between the second rotating shaft 120 and the inner air bearing 510, so that turbine cooling air supplied from the air supply passage 511 of the inner air bearing 510 to the shaft inner bearing cavity 122 does not leak as much as possible to ensure a cooling effect.

In some alternative embodiments, the first air aperture 512 is a stepped aperture with a large aperture end facing the inner air bearing 510, which may further increase the incoming air flow velocity to further enhance the air film strength.

In some alternative embodiments, the first air holes 512 are arranged in more than two circles along the axial direction to increase the air output, ensure the strength of the air film, and increase the stability of the inner air bearing 510.

In some alternative embodiments, as shown in fig. 7, fig. 7 is a partial schematic view of another embodiment of the inner air bearing in fig. 1, the air supply passage 511 has a flared opening 513 with an increased diameter corresponding to the air outlet, and the first air hole 512 is opened corresponding to the flared opening 513. The provision of a flared mouth 513 of a larger diameter than the original gas supply passage 511 facilitates a machining tool (e.g., a hook cutter) to machine the large-diameter end of the first gas orifice 512 deep into the flared mouth 513, which may then be used to machine the small-diameter end of the first gas orifice 512, for example, by laser drilling.

Further, the pressure of the gas supplied through the gas supply passage 511 is higher than the turbine front pressure of the third turbine 330, which ensures that the cooling gas can be ejected through the turbine film hole 332 against the turbine front pressure to form a cooling film.

Specifically, an air inlet of the air supply passage 511 is connected to an exhaust end of the second compressor 220, and the pressure at the exhaust end of the second compressor 220 is higher than the pre-turbine pressure of the third turbine 330.

Further, as shown in FIG. 1, the gas turbine also includes a first vane assembly 710, the first vane assembly 710 being located between the third turbine 330 and the exhaust end of the combustor 400 and disposed proximate to the third turbine 330. The stationary blade in the first stationary blade assembly 710 is a fluid cooling blade with a hollow structure to improve the cooling effect of the first stationary blade assembly 710, which is beneficial to improving the temperature of the exhaust end of the combustion chamber 400. The fluid cooling blade can be an air film cooling blade or a liquid film cooling blade, or a blade with a gas or liquid circulation cooling channel inside.

Further, as shown in FIG. 8, FIG. 8 is a partial schematic view of yet another embodiment of the third turbine of FIG. 1, the first vane assembly 710 includes a first support 711, a second support 713, and a blade set 712 having an axial air passage between the two supports. The turbine 330 further includes a turbine thrust nozzle hole 335, an air outlet of the turbine thrust nozzle hole 335 faces a wall surface of the first seat 711 of the first guide vane assembly 710, and the pressure air ejected from the turbine thrust nozzle hole 335 can form an axial thrust force, which is beneficial to axial balance of a rotor system of the gas turbine. Further, air sealing holes 336 may be disposed on two sides of the turbine thrust nozzle hole 335, a diameter of the air sealing holes 336 is smaller than a diameter of the turbine thrust nozzle hole 335, and the pressure air ejected from the air sealing holes 336 may form an air curtain seal to seal the air ejected from the turbine thrust nozzle hole 335 as much as possible to improve the thrust.

Further, as shown in fig. 9, fig. 9 is a schematic structural diagram of a gas turbine according to another embodiment of the present invention, and the second turbine 320 may also be a film-cooled turbine having a hollow structure. The specific structure thereof may be substantially the same as that of the third turbine 330.

Further, as shown in fig. 9, the gas turbine further includes a second vane assembly 720, the second vane assembly 720 is located between the third turbine 330 and the second turbine 320 and is disposed near the second turbine 320, and the vanes in the second vane assembly 720 are fluid-cooled blades having a hollow structure. The specific structure of the second vane assembly 720 may be substantially the same as the first vane assembly 710.

Further, as shown in FIG. 1, the gas turbine also includes a first electric machine 910 and a second electric machine 920. The first motor 910 is mounted on the first rotating shaft 110 and is located between the first compressor 210 and the first turbine 310 or located on a side of the first compressor 210 facing away from the first turbine 310. Since the high temperature between the second compressor 220 and the third turbine 330 may cause the efficiency of the motor to be reduced, the second motor 920 may be mounted on the second rotating shaft 120 and located on the side of the second compressor 220 away from the third turbine 330.

The power of the first motor 910 is greater than or equal to the power of the second motor 920. It should be understood that the dimensions and proportions of the first motor 910 and the second motor 920 shown in the figures are not representative of their actual dimensions and proportions, nor are they indicative of the magnitude and comparison of power.

Further, the first motor 910 and/or the second motor 920 may be superconducting motors, which can reduce the volume of the motors and achieve higher motor efficiency (even 100%). And the superconducting motor can provide larger power (hundreds of kilowatts or megawatts), and can be applied to a gas turbine with larger power. In this embodiment, the fuel of the gas turbine is reused as the coolant of the superconducting electrical machine, and the fuel may be liquefied hydrogen, liquefied natural gas, or the like.

Further, as shown in fig. 10 and 11, fig. 10 is a schematic structural view of a gas turbine according to still another embodiment of the present invention, and fig. 11 is a schematic structural view of an embodiment of the second rotating shaft 120 in fig. 10, and the second rotating shaft 120 includes a first shaft section 121 and a second shaft section 125 which are arranged in this order along the axial direction. The first shaft section 121 and the second shaft section 125 are connected by a coupling 128, and the coupling 128 has a gas seal cavity 129 communicating with gas passages in the first shaft section 121 and the second shaft section 125. Wherein the second motor 920 is mounted to the second shaft segment 125. In embodiments where the gas turbine includes an inner air bearing 510, the inner air bearing 510 is located at an end of the second shaft segment 125 axially distal from the first shaft segment 121. The length of the second rotating shaft 120 can be prevented from being too long, the rotor dynamics parameters are difficult to process and teach, and after the first shaft section 121 and the second shaft section 125 which are connected through the coupler 128 are arranged, the size and the power of the second motor 920 can be increased as required, so that the power balance and allocation between the first motor 910 and the second motor 920 are met.

Specifically, as shown in FIG. 11, the proportions of the first shaft segment 121 and the second shaft segment 125 in FIG. 11 are adjusted as compared to FIG. 10 for ease of illustration. The first shaft section 121 has a shaft coupling through hole 123 and a shaft inner air passage 124, and the second shaft section 125 has a shaft inner air passage 126 and a shaft inner bearing cavity 122. The coupling 128 is connected to the first shaft section 121 and the second shaft section 125 in an airtight manner and has an air seal cavity 129, and the air seal cavity 129 forms an in-shaft air path of the second rotating shaft 120 with the in-shaft air path 124 and the in-shaft air path 126.

Further, as shown in fig. 12, fig. 12 is a schematic structural view of a gas turbine according to still another embodiment of the present invention, one end of the second rotating shaft 120, at which the inner air bearing 510 is disposed, has a nose cone 126, and the inner air bearing 510 has a nose cone air supply hole. The nose cone 126 may be integrally formed on the second shaft 120, or may be fixedly mounted on the second shaft 120 by welding or connecting means. The end of nose cone 126 axially toward inner air bearing 510 has a thrust face. Inner air bearing 510 has a nose cone air supply hole, the air inlet end of which may communicate with air supply passage 511, the air outlet of which faces the thrust face of nose cone 126. The pressure gas sprayed from the air supply hole of the nose cone can provide axial thrust to the nose cone 126, so that the axial force balance of the rotor system is facilitated. In other embodiments, the air inlet end of the nose cone air supply hole can be connected to the air source through other channels. In some embodiments, the side of nose cone 126 distal from inner air bearing 510 may be provided with a thrust bearing.

Further, a thrust disc may be disposed on the first rotating shaft 110 and/or the second rotating shaft 120, and thrust bearings may be disposed on both axial sides of the thrust disc.

Further, as shown in fig. 1,9,10, and 12, the gas turbine further includes a plurality of first radial bearings 520, and the first radial bearings 520 are outboard radial air bearings sleeved on the outer circumferential surfaces of the first rotating shaft 110 and/or the second rotating shaft 120. The position of the first radial bearing 520 may be one or more of the positions shown in the above figures, or may be other axial positions not shown.

Further, as shown in fig. 13, fig. 13 is a schematic structural view of a gas turbine according to still another embodiment of the present invention, and fig. 13 shows an example in which the inner air bearing 510 is located at an end of the second rotating shaft 120 axially close to the first rotating shaft 110.

Further, as shown in fig. 14, fig. 14 is a schematic structural diagram of a gas turbine according to still another embodiment of the present invention, and the gas turbine further includes a heat regenerator HR to further improve the efficiency of the gas turbine. Specifically, the regenerator HR includes a first passage and a second passage that can exchange heat, an exhaust end of the second compressor 220 communicates with an intake end of the first passage, an intake end of the combustion chamber 400 communicates with an exhaust end of the first passage, an exhaust end of the first turbine 310 communicates with an intake end of the second passage, and an exhaust end of the second passage communicates with the outside atmosphere or other external components.

In the description of the present invention, it is to be understood that the terms "central," "longitudinal," "lateral," "length," "width," "thickness," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "clockwise," "counterclockwise," "axial," "radial," "circumferential," and the like are used in the orientations and positional relationships indicated in the drawings for convenience in describing the invention and to simplify the description, but are not intended to indicate or imply that the device or element so referred to must have a particular orientation, be constructed in a particular orientation, and be operated in a particular manner, and are not to be construed as limiting the invention. Furthermore, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of the present invention, "a plurality" means two or more unless otherwise specified.

In the description of the present invention, it should be noted that, unless otherwise explicitly specified or limited, the terms "mounted," "connected," and "connected" are to be construed broadly and may be, for example, fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meanings of the above terms in the present invention can be understood in specific cases to those skilled in the art.

In the description of the present invention, reference to the description of the terms "one embodiment," "some embodiments," "an illustrative embodiment," "an example," "a specific example," or "some examples" or the like means that a particular feature, structure, material, or characteristic described in connection with the embodiment or example is included in at least one embodiment or example of the present invention. In this specification, the schematic representations of the terms used above do not necessarily refer to the same embodiment or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments or examples.

While the invention has been described with reference to the above embodiments, these embodiments are not intended to be exhaustive or to limit the invention to the precise embodiments disclosed. Obviously, many modifications and variations are possible in light of the above teaching. The embodiments were chosen and described in order to best explain the principles of the invention and the practical application, to thereby enable others skilled in the art to best utilize the invention and various embodiments with various modifications as are suited to the particular use contemplated. The invention is limited only by the claims and their full scope and equivalents.

Claims

1. The utility model provides a gas turbine, its characterized in that includes first pivot and the second pivot that arranges in proper order along the axial to and the combustion chamber, keeping away from along the axial of first pivot second pivot one end is installed first compressor and is close to along the axial first turbine is installed to second pivot one end, keeping away from along the axial of second pivot first pivot one end is installed the second compressor and is close to along the axial second turbine and third turbine are installed in proper order to first pivot one end, the inlet end of second compressor with the exhaust end intercommunication of first compressor, the exhaust end of second compressor with the inlet end intercommunication of combustion chamber, the exhaust end of combustion chamber with the inlet end intercommunication of third turbine, the gas film cooling turbine of third turbine for having hollow structure.

2. The gas turbine of claim 1, wherein at least one axial end of said second shaft is provided with an at least partially built-in inner air bearing having an air supply passage for supplying cooling air to said third turbine through the interior of said second shaft.

3. The gas turbine of claim 2, wherein said second shaft has a nose cone at an end thereof where said inner air bearing is located, said inner air bearing having a nose cone air supply hole; and/or the presence of a gas in the atmosphere,

the gas turbine further comprises a plurality of first radial bearings, and the first radial bearings are external radial air bearings sleeved on the peripheral surfaces of the first rotating shaft and/or the second rotating shaft.

4. The gas turbine according to claim 1, wherein the second turbine is located between the first turbine and the third turbine in an axial direction, and the gas from the exhaust end of the combustor is discharged after passing through the third turbine, the second turbine, and the first turbine in this order; and/or the presence of a gas in the atmosphere,

the gas turbine further includes a first vane assembly located between the third turbine and the exhaust end of the combustor and disposed proximate to the third turbine, vanes in the first vane assembly being fluid cooled blades having a hollow structure.

5. The gas turbine according to claim 1, wherein the second turbine is a film-cooled turbine having a hollow structure; and/or the presence of a gas in the gas,

the gas turbine further includes a second vane assembly located between the third turbine and the second turbine and disposed proximate to the second turbine, vanes in the second vane assembly being fluid-cooled blades having a hollow structure.

6. A gas turbine according to claim 1, further comprising a first electric machine and a second electric machine,

the first motor is arranged on the first rotating shaft and is positioned between the first compressor and the first turbine or positioned on one side of the first compressor, which is far away from the first turbine;

the second motor is arranged on the second rotating shaft and is positioned on one side, far away from the third turbine, of the second compressor;

the power of the first motor is greater than or equal to the power of the second motor.

7. The gas turbine of claim 6, wherein the second shaft includes a first shaft section and a second shaft section arranged axially in sequence, the first shaft section and the second shaft section being connected by a coupling having a gas seal cavity communicating with gas passages in the first shaft section and the second shaft section;

the second motor is mounted on the second shaft section;

the gas turbine comprises an inner air bearing, and the inner air bearing is positioned at one end, far away from the first shaft section, of the second shaft section in the axial direction.

8. A gas turbine according to claim 6, wherein the first and/or second electrical machine is a superconducting electrical machine, the fuel of the gas turbine being reused as a coolant for the superconducting electrical machine.

9. The gas turbine according to claim 1,

the pressure ratio of the second compressor is more than or equal to that of the first compressor;

the second compressor is a centrifugal compressor;

the first compressor is selected from one or more of a centrifugal compressor, an axial compressor and a diagonal compressor.

10. The gas turbine according to claim 1, further comprising a regenerator HR, the regenerator HR including a first passage and a second passage which are heat-exchangeable, an exhaust end of the second compressor communicating with an intake end of the first passage, an intake end of the combustion chamber communicating with an exhaust end of the first passage, an exhaust end of the first turbine communicating with an intake end of the second passage, and an exhaust end of the second passage communicating with an external atmosphere or other external components.